Air temperature sensor

ABSTRACT

A total air temperature sensor can include a first airfoil having a heated first surface, a second airfoil having a second surface spaced from the first surface and defining a sensor chamber, a temperature sensor located within the chamber, and a sheath surrounding the temperature sensor.

BACKGROUND OF THE INVENTION

Turbine engines, and particularly gas or combustion turbine engines, arerotary engines that extract energy from a flow of combusted gasespassing through the engine onto a multitude of rotating turbine blades.Gas turbine engines have been used for land and nautical locomotion andpower generation, but are most commonly used for aeronauticalapplications such as airplanes or helicopters. In airplanes, gas turbineengines are used for propulsion of the aircraft.

During operation of a turbine engine, the total air temperature alsoknown as stagnation temperature can be measured by a specially designedtemperature probe mounted on the surface of the aircraft or the interiorwalls of the turbine engine. The probe is designed to bring the air torest relative to the aircraft. The air experiences an adiabatic increasein temperature as it is brought to rest and measured, and the total airtemperature is therefore higher than the ambient air temperature. Totalair temperature is an essential input for calculating static airtemperature and true airspeed. Total air temperature sensors can beexposed to adverse conditions including high Mach numbers and icingconditions, as well as water and debris, which may affect the readingprovided by the sensor.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, the disclosure relates to an air temperature sensorsuitable for use on an aircraft, the temperature sensor comprising ahousing defining an interior and having at least a portion with anairfoil cross section to define an airfoil portion with an upper surfaceand a lower surface, a temperature sensor located within the airfoilportion, an airflow path having an inlet in the upper surface of thehousing and extending through the housing to the temperature sensor toprovide for diverted air from air flowing along the upper surface tocontact the temperature sensor; and a set of fluid passageways definedwithin the interior and having an inlet and a set of outlets locatedwithin the housing and where the set of fluid passageways are configuredto receive hot bleed air via the inlet and disperse the hot bleed air tothe set of outlets to heat at least a portion of the airfoil portion.

In another aspect, the disclosure relates to an air temperature sensor,comprising a housing having a skin and defining an interior, atemperature sensor having a first portion located within the interiorand a second portion extending through a portion of the housing and atleast partially adjacent a portion of the skin, and a set of fluidpassageways defined within the interior and configured to receive hotbleed air and disperse the hot bleed air to at least two separateportions of the skin.

In yet another aspect, the disclosure relates to a method of forming atotal air temperature sensor housing, the method comprising forming, viaadditive manufacturing a housing having an exterior surface and definingan interior and having at least a portion with an airfoil cross sectionto define an airfoil portion with an upper surface and a lower surfaceand having a set of fluid passageways defined within the interior andhaving an inlet and a set of outlets located within the housing andwhere the set of fluid passageways are configured to receive hot bleedair via the inlet and disperse the hot bleed air to the set of outletsto heat at least a portion of the exterior surface of at least a portionof the airfoil cross section.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic cross-sectional diagram of a turbine engine for anaircraft with a total air temperature sensor.

FIG. 2 is an enlarged isometric view of the total air temperature sensorin a partially cut-away portion of the engine of FIG. 1

FIG. 3 is an exploded view of the total air temperature sensor of FIG.2.

FIG. 4 is a cross-sectional view of the total air temperature sensortaken along line IV-IV of FIG. 2.

FIG. 5 is a cross-sectional view of the total air temperature sensortaken along line V-V of FIG. 4.

FIG. 6 is an enlarged partial cross-sectional view of a portion of thetotal air temperature sensor of FIG. 4.

FIG. 7 is a partial cross-sectional view of the total air temperaturefrom FIG. 2 with a dispersion chamber.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

The described embodiments of the present disclosure are directed to anair temperature sensor for an aircraft turbine engine. It will beunderstood, however, that the disclosure is not so limited and may havegeneral applicability within an engine, as well as in non-aircraftapplications, such as other mobile applications and non-mobileindustrial, commercial, and residential applications.

As used herein, the term “forward” or “upstream” refers to moving in adirection toward the engine inlet, or a component being relativelycloser to the engine inlet as compared to another component. The term“aft” or “downstream” used in conjunction with “forward” or “upstream”refers to a direction toward the rear or outlet of the engine or beingrelatively closer to the engine outlet as compared to another component.

Additionally, as used herein, the terms “radial” or “radially” refer toa dimension extending between a center longitudinal axis of the engineand an outer engine circumference. A “set” as used herein can includeany number of a particular element, including only one.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, forward, aft, etc.) are only used foridentification purposes to aid the reader's understanding of the presentdisclosure, and do not create limitations, particularly as to theposition, orientation, or use of the disclosure. Connection references(e.g., attached, coupled, connected, and joined) are to be construedbroadly and can include intermediate members between a collection ofelements and relative movement between elements unless otherwiseindicated. As such, connection references do not necessarily infer thattwo elements are directly connected and in fixed relation to oneanother. The exemplary drawings are for purposes of illustration onlyand the dimensions, positions, order, and relative sizes reflected inthe drawings attached hereto can vary.

FIG. 1 is a schematic cross-sectional diagram of a gas turbine engine 10for an aircraft. The engine 10 has a generally longitudinally extendingaxis or centerline 12 extending forward 14 to aft 16. The engine 10includes, in downstream serial flow relationship, a fan section 18including a fan 20, a compressor section 22 including a booster or lowpressure (LP) compressor 24 and a high pressure (HP) compressor 26, acombustion section 28 including a combustor 30, a turbine section 32including a HP turbine 34, and a LP turbine 36, and an exhaust section38.

The fan section 18 includes a fan casing 40 surrounding the fan 20. Thefan 20 includes a plurality of fan blades 42 disposed radially about thecenterline 12. The HP compressor 26, the combustor 30, and the HPturbine 34 form a core 44 of the engine 10, which generates combustiongases. The core 44 is surrounded by core casing 46, which can be coupledwith the fan casing 40. A total air temperature (TAT) sensor 90 can bedisposed in the fan casing 40 as shown; however, this example is notmeant to be limiting and the TAT sensor 90 may be positioned in otherlocations in the turbine engine 10.

A HP shaft or spool 48 disposed coaxially about the centerline 12 of theengine 10 drivingly connects the HP turbine 34 to the HP compressor 26.A LP shaft or spool 50, which is disposed coaxially about the centerline12 of the engine 10 within the larger diameter annular HP spool 48,drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.The spools 48, 50 are rotatable about the engine centerline and coupleto a plurality of rotatable elements, which can collectively define arotor 51.

The LP compressor 24 and the HP compressor 26 respectively include aplurality of compressor stages 52, 54, in which a set of compressorblades 56, 58 rotate relative to a corresponding set of staticcompressor vanes 60, 62 (also called a nozzle) to compress or pressurizethe stream of fluid passing through the stage. In a single compressorstage 52, 54, multiple compressor blades 56, 58 can be provided in aring and can extend radially outwardly relative to the centerline 12,from a blade platform to a blade tip, while the corresponding staticcompressor vanes 60, 62 are positioned upstream of and adjacent to therotating blades 56, 58. It is noted that the number of blades, vanes,and compressor stages shown in FIG. 1 were selected for illustrativepurposes only, and that other numbers are possible.

The blades 56, 58 for a stage of the compressor can be mounted to a disk61, which is mounted to the corresponding one of the HP and LP spools48, 50, with each stage having its own disk 61. The vanes 60, 62 for astage of the compressor can be mounted to the core casing 46 in acircumferential arrangement.

The HP turbine 34 and the LP turbine 36 respectively include a pluralityof turbine stages 64, 66, in which a set of turbine blades 68, 70 arerotated relative to a corresponding set of static turbine vanes 72, 74(also called a nozzle) to extract energy from the stream of fluidpassing through the stage. In a single turbine stage 64, 66, multipleturbine blades 68, 70 can be provided in a ring and can extend radiallyoutwardly relative to the centerline 12 while the corresponding staticturbine vanes 72, 74 are positioned upstream of and adjacent to therotating blades 68, 70. It is noted that the number of blades, vanes,and turbine stages shown in FIG. 1 were selected for illustrativepurposes only, and that other numbers are possible.

The blades 68, 70 for a stage of the turbine can be mounted to a disk71, which is mounted to the corresponding one of the HP and LP spools48, 50, with each stage having a dedicated disk 71. The vanes 72, 74 fora stage of the compressor can be mounted to the core casing 46 in acircumferential arrangement.

Complementary to the rotor portion, the stationary portions of theengine 10, such as the static vanes 60, 62, 72, 74 among the compressorand turbine section 22, 32 are also referred to individually orcollectively as a stator 63. As such, the stator 63 can refer to thecombination of non-rotating elements throughout the engine 10.

In operation, the airflow exiting the fan section 18 is split such thata portion of the airflow is channeled into the LP compressor 24, whichthen supplies pressurized air 76 to the HP compressor 26, which furtherpressurizes the air. The pressurized air 76 from the HP compressor 26 ismixed with fuel in the combustor 30 and ignited, thereby generatingcombustion gases. Some work is extracted from these gases by the HPturbine 34, which drives the HP compressor 26. The combustion gases aredischarged into the LP turbine 36, which extracts additional work todrive the LP compressor 24, and the exhaust gas is ultimately dischargedfrom the engine 10 via the exhaust section 38. The driving of the LPturbine 36 drives the LP spool 50 to rotate the fan 20 and the LPcompressor 24.

A portion of the pressurized airflow 76 can be drawn from the compressorsection 22 as bleed air 77. The bleed air 77 can be drawn from thepressurized airflow 76 and provided to engine components requiringcooling. The temperature of pressurized airflow 76 entering thecombustor 30 is significantly increased. As such, cooling provided bythe bleed air 77 is necessary for operating of such engine components inthe heightened temperature environments.

A remaining portion of the airflow 78 bypasses the LP compressor 24 andengine core 44 and exits the engine assembly 10 through a stationaryvane row, and more particularly an outlet guide vane assembly 80,comprising a plurality of airfoil guide vanes 82, at the fan exhaustside 84. More specifically, a circumferential row of radially extendingairfoil guide vanes 82 are utilized adjacent the fan section 18 to exertsome directional control of the airflow 78.

Some of the air supplied by the fan 20 can bypass the engine core 44 andbe used for cooling of portions, especially hot portions, of the engine10, and/or used to cool or power other aspects of the aircraft. In thecontext of a turbine engine, the hot portions of the engine are normallydownstream of the combustor 30, especially the turbine section 32, withthe HP turbine 34 being the hottest portion as it is directly downstreamof the combustion section 28. Other sources of cooling fluid can be, butare not limited to, fluid discharged from the LP compressor 24 or the HPcompressor 26.

FIG. 2 more clearly depicts the TAT sensor 90 in a cut away portion ofthe engine 10. A mounting section 92 having a suitable mounting portion94 can be included in the TAT sensor 90. A wiring housing 96 can beincluded in the mounting section 92 and can be coupled to an electricalconduit 98. The mounting section 92 can be any suitable mounting portion94 and is not meant to be limiting. A housing 102 is mounted at an uppersection 104 of the housing 102 to a portion of the aircraft engine 10 atthe mounting section 92. A tube inlet 108 couples to the housing 102 andis coupled to a source of hot bleed air. By way of non-limiting examplebleed air 110 is illustrated as entering the tube inlet 108.

A skin 100 defines an exterior surface 103 of the housing 102 of the TATsensor 90. At least one set of outlets 101 is included in the skin 100.The skin 100 can include at least two separate portions of the skin 100a, 100 b that can be wetted surfaces. A wetted surface can be anysurface susceptible to condensation and ice accumulation.

A lower section 112 of the housing 102 defines an airfoil portion 114. Aportion of the skin 100 can form the airfoil portion 114 of the lowersection 112. The airfoil portion 114 can have a concave side, or anupper surface 116 and a convex side, or a lower surface 118. The airfoilportion 114 can extend from a leading edge 115 to a trailing edge 117. Atemperature sensor inlet 120 in the upper surface 116 extends throughthe portion of the skin 100 b to an outlet 122 (FIG. 3) to provide adiverted airflow path (DAP) for a portion of the pressurized airflow 76.

Turning to FIG. 3, an exploded view of the TAT sensor 90 is illustrated.The TAT sensor 90 is illustrated in a different orientation than that ofFIG. 2 to more clearly show the temperature sensor outlet 122 adjacentan open portion 124 defined by the lower section 112 of the housing. Theopen portion 124 defined by the housing 102 separates the two portionsof the skin 100 a, 100 b to define the temperature sensor outlet 122there between. The temperature sensor outlet 122 is proximate thetrailing edge 117 and on the lower surface 118 of the airfoil portion114.

A tube, by way of non-limiting example a piccolo tube 132 extends from afirst end 134 to a second end 136. The first end 134 is coupled to thetube inlet 108 and the second end 136 can extend into the housing 102.

A temperature sensor assembly 139 includes an upper sheath 140,protective sleeving 142, and a temperature sensor 144. The temperaturesensor 144 is a total air temperature sensor suitable for use on anaircraft, within the engine 10.

The temperature sensor assembly 139 can further include a lockingmechanism 148 and a lower sheath 150. The locking mechanism 148 can belocated within the housing 102. The lower sheath 150 can include slotopening 151 through which diverted air along the diverted airflow path(DAP) can contact the temperature sensor 144. The locking mechanism 148can be shaped in any suitable manner and orientated in any suitablemanner with respect to the diverted airflow path (DAP) and thetemperature sensor 144. At least one rib 126 with an aperture 128 can belocated within the open portion 124. When assembled, the at least onerib 126 can aide in stabilizing the lower sheath 150 surrounding thetemperature sensor 144.

More specifically, when assembled, as in FIG. 4, the lower sheath 150 islocated within the open portion 124 defined by the housing 124. Thelower sheath 150 extends through the aperture 128 of the at least onerib 126. The locking mechanism 148 of the temperature sensor assembly139 encompasses the protective sleeving 142 and upper sheath 140 of thetemperature sensor 144. The lower sheath 150 encompasses the temperaturesensor 144.

An interior 158 of the housing 124 is defined at least in part by theskin 100. A first portion 158 a of the interior 158 can be includedwithin the first portion of the skin 100 a. A second portion 158 b ofthe interior 158 can be located within the second portion of the skin100 b.

A dispersion chamber 166 is located within the interior 158 of thehousing 124. The dispersion chamber 166 can be defined by a set of walls192 and be fluidly coupled to a transfer tube 182 and a set ofintermediate conduits 198.

An inlet 162 into the dispersion chamber 166 is defined by a tip 163having a set of spray openings 164. The tip 163 defines the inlet 162and is operably coupled to one of the set of walls 192. The second end136 of the piccolo tube 132 is coupled to the interior 158 of thehousing 124 via the tip 163.

The set of outlets 101 can include multiple sets of outlets 101 providedwithin the skin 100. A first set of outlets 101 a is provided within thefirst portion 158 a and a second set of outlets 101 b is provided withinthe second portion 158 b.

A set of fluid passageways 172 is provided throughout the interior 158.The set of fluid passageways 172 can include a first fluid passageway172 a within the first portion 158 a of the interior 158. The firstfluid passage way 172 a includes a first set of channels 174 a fluidlyconnecting the dispersion chamber 166 to the first set of outlets 101 a.The exemplary first set of channels 174 a includes parallel channels 174a of similar width and length coupled by a first turn 176. In thismanner the first fluid passageway 172 a doubles back along a portion 178of a length (L) of the housing 102. It is contemplated that the firstset of channels 174 a are oriented in any suitable manner including butnot limited to in parallel, in serpentine, or in series patterns, suchthat the dispersion chamber 166 is fluidly coupled to the first set ofoutlets 101 a within the first portion 158 a of the interior 158.

A second fluid passageway 172 b within the second portion 158 b of theinterior 158 includes a second set of channels 174 b fluidly connectingthe dispersion chamber 166 to the second set of outlets 101 b.

A set of dead air spaces 160 can also be included within the housing102. The set of dead air spaces 160 are fluidly separate from the set offluid passageways 172. By way of non-limiting example, the dead airspaces 160 can be located within the housing 102 where hot air need notbe dispersed. The set of dead air spaces 160 can be located between thefirst portion 158 a and the second portion 158 b, such that at least aportion of the set of dead air spaces 160 extends parallel to the firstand second set of channels 174 a, 174 b.

FIG. 5 more clearly shows a portion of the dispersion chamber 166 andthe set of spray openings 164. A cap 191 forms the tip 163 and the setof spray openings 164 are located about portions of the cap 191. The cap191, can be rounded having a perimeter 190 and the set of spray openings164 can be spaced about the perimeter 190. At least one of the sprayopenings 164 a can be formed at a distal end 196 of the tip 163. Asillustrated the set of spray openings 164 can be a plurality of sprayopenings 164 equally spaced about the perimeter 190 and configured tospray the hot bleed air 110 within the dispersion chamber 166.

The set of walls 192 forming the dispersion chamber 166 can include anangled surface 192 a, a series surface 192 b, a parallel surface 192 c,and an inlet surface 192 d. A set of corners 194 can be defined whereany two of the set of walls meet.

The set of spray openings 164 is configured to direct hot bleed air intothe dispersion chamber 166, onto the walls forming the dispersionchamber 166, and into the fluid passageways 172. In the illustratedexample, a first portion 110 a of the hot bleed air is directed into thefirst fluid passageways 172 a via the set of intermediate conduits 198.The set of spray openings is further configured to direct a secondportion 110 b of hot bleed air into the second fluid passageways 172 b(FIG. 4) via the transfer tube 182. The hot bleed air 110 can beseparated into further portions of hot bleed air 110 c, wherein at leastone of the further portions of hot bleed air 110 c is introduced to theset of corners 194 and/or the set of walls 192. In particular at leastone spray opening 164 b is oriented such that it heats the angledsurface 192 a.

Turning to FIG. 6, the cross section through a portion of the airfoilformed by the lower section 112 of the TAT sensor 90 more clearlyillustrates a portion of the set of fluid passageways 172. It can beseen that the first and second fluid passageways 172 a, 172 b arelocated on opposite sides of the housing 120 and on either side of thediverted airflow path (DAP). It can also be seen that an airfoil crosssection 154 can be asymmetrical although this need not be the case.

Additionally, it is also more clearly depicted, that the second set ofchannels 174 b can be oriented in any suitable manner including, but notlimited to, in parallel. Further, it can be seen that the set ofchannels 174 need not have the same shape or cross-sectional area. Thesecond set of channels 174 b can also include a second turn 180illustrated in phantom. In this manner the second fluid passageway 172 bdoubles back similarly to the first fluid passageway 172 a. It isfurther contemplated that the second set of channels 174 b can be in anyorientation including in serpentine, or in series patterns, and have anyvarying volumes such that the inlet 162 is fluidly coupled to the secondset of outlets 101 b within the second portion 158 b of the interior158.

The set of dead air spaces 160 is proximate the temperature sensor 144.In this manner, the lower sheath 150 along with the set of dead airspaces 160 together can shield the temperature sensor 144 from heatwithin the first and second set of channels 174 a, 174 b.

During operation the diverted airflow path (DAP) flows through thetemperature sensor inlet 120 and over the lower sheath 150. Thetemperature sensor 144 is exposed in such a manner as to record atemperature of the diverted airflow path (DAP). During operation theexterior surface 103 of the airfoil portion 114 may become warmed fromheat within the first and second set of channels 174 a, 174 b. The lowersheath 150 channels any warmed air at the exterior surface 103 away fromthe temperature sensor 144 and prevents the warmed air from reaching thetemperature sensor 144 reducing deicing errors. The diverted airflowpath (DAP) and lower sheath 150 function to form an airflow stagnationarea about the temperature sensor 144 to provide for a total airtemperature reading by the temperature sensor 144.

FIG. 7 illustrates a plurality of airflow paths shown in a partialcross-section of the housing 102. The airflow paths within the housing102 are defined at least in part by the set of fluid passageways 172.

During operation, the hot bleed air 110 can enter at the inlet 162 andbe dispersed by the set of spray openings 164 within the dispersionchamber 166. A first portion 110 a of the hot bleed air 110 flowsthrough the intermediate conduits 198 and along the first fluidpassageways 172 a defining a first hot airflow path (FAP). The first hotairflow path (FAP) can flow along the length (L) of the housing 102 andat least in part along the leading edge 115 of the airfoil portion 114.The first hot airflow path (FAP) can turn at the first turn 176, andexit through the first set of outlets 101 a.

The transfer tube 182 changes from an orientation perpendicular thelength L to an orientation parallel the length L at a third turn 184.The transfer tube 182 fluidly couples the inlet 162 to the second fluidpassageways 172 b. A second portion 110 b of the hot bleed air 110enters at the inlet 162 and flows along the second fluid passageways 172b. The second hot airflow path (SAP) flows through the transfer tube 182perpendicular to the length (L) of the housing, turns at the third turn184 to flow along the portion 178 of the housing 102, turns again at thesecond turn 180 and exits through the second set of outlets 101 b.

The first hot airflow path (FAP) is configured to heat the portion ofthe skin 100 a proximate the leading edge 115 of the airfoil portion114. The second hot airflow path (SAP) is configured to heat the skin100 b proximate the open portion 124 of the airfoil portion 114.Together the first hot airflow path (FAP) and the second hot airflowpath (SAP) heat the exterior 103 surface of the housing 102 to preventice buildup along the airfoil portion 114.

A method of forming the TAT sensor 90 as described herein can includeforming, via additive manufacturing the housing 102 with the skin 100defining the interior 158 and including the airfoil cross section 154defining the airfoil portion 114. The additive manufacturing can formthe airfoil portion including the upper surface 116 and the lowersurface 118. The additive manufacturing can form the set of fluidpassageways 172 within the interior 158 and the inlet 162 and the set ofoutlets 101 located within the housing. The additive manufacturing canform the cap 191 integrally with a remainder of the housing 102; thus,forming the tip 163 and the set of spray openings 164. The additivemanufacturing is done such that the set of fluid passageways 172 areconfigured to receive the hot bleed air 110 via the inlet 162 anddisperse the hot bleed air 110 to the set of outlets 101 to heat atleast a portion of the exterior surface. The additive manufacturing byway of non-limiting examples, can include direct metal laser melting ordirect metal laser sintering.

Benefits associated with the disclosure discussed herein includespneumatically supplying heated air and directing the heated air tocritical locations of the sensor housing without impacting the sensor'sreading. Location and size of the channels can be optimized usingadditive manufacturing without relying on current conventionalsubtractive manufacturing, by way of non-limiting example machining,drilling, and grinding.

Typical sensor exposed to an icing environment have been mechanicallydesigned or positioned in the environment so that any large amounts ofice shed from the sensor will not damage items behind it. This limitslocation selection for the TAT sensor, and therefore limits theperformance of the TAT sensor. Eliminating the ice shedding with heatingsystems within the TAT sensor improves location possibilities.Additionally considering increased sensitivity to ice shedding incurrent engine design, TAT sensors with minimal to no ice shedding arepreferred.

Additively manufacturing the TAT sensor allows for locating of theheating channels along any desired location. Assembly time of the TATsensor is also reduced due to the housing being additively manufactured.

Additionally the dispersion chamber as described herein utilizes anoutlet with a set of spray openings for direct heating to areas of theTAT sensor with high ice concentrations. The diffused hot air is thentransferred to the airfoil portion of the TAT sensor to further removeice buildup.

It should be understood that application of the disclosed design is notlimited to turbine engines with fan and booster sections, but isapplicable to turbojets and turbo engines as well.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. An air temperature sensor, comprising: a housinghaving an upper section configured to be coupled with a portion of anaircraft engine and a lower section having an airfoil cross section; adispersion chamber located within the housing; a set of fluidpassageways defined within the housing and fluidly coupled to thedispersion chamber; a piccolo tube having a first end fluidly coupled toreceive bleed air from a portion of the aircraft engine and a secondend, fluidly coupled to the first end, and where the second end includesa set of spray openings configured to allow bleed air to spray intoportions of the dispersion chamber and the set of fluid passageways. 2.The air temperature sensor of claim 1 wherein the set of spray openingsare spaced about a perimeter of the second end.
 3. The air temperaturesensor of claim 2 wherein the set of spray openings are equally spacedabout a perimeter of the second end.
 4. The air temperature sensor ofclaim 1 wherein the second end comprises a tip with the set of sprayopenings located therein.
 5. The air temperature sensor of claim 4wherein the tip is located within the dispersion chamber and configuredto direct the bleed air to surfaces forming the dispersion chamber. 6.The air temperature sensor of claim 5 wherein the set of fluidpassageways includes a first fluid passageway within a first portion ofthe housing and a second fluid passageway with a second portion of thehousing.
 7. The air temperature sensor of claim 6 wherein the set ofspray openings are configured to direct the bleed air into the firstfluid passageway and the second fluid passageway.
 8. The air temperaturesensor of claim 5 wherein the set of spray openings are configured tospray the bleed air into at least one corner of the dispersion chamber.9. The air temperature sensor housing of claim 4 wherein at least someof the set of spray openings are staggered about a circumference of tip.10. The air temperature sensor of claim 9 wherein at least one of theset of spray openings is formed at a distal end of the tip.
 11. The airtemperature sensor of claim 1, further comprising an intermediateconduit fluidly coupling the dispersion chamber to at least one of theset of fluid passageways.
 12. The air temperature sensor of claim 11wherein at least one of the set of spray openings is configured todirect the bleed air into the intermediate conduit.
 13. An airtemperature sensor, comprising: a housing having a skin and defining aninterior; a temperature sensor having a first portion located within theinterior and a second portion extending through a portion of the housingand at least partially adjacent a portion of the skin; a set of fluidpassageways defined within the interior; and a tube having a first endfluidly coupled to receive bleed air from a portion of an aircraftengine and a second end, fluidly coupled to the first end, and locatedwithin the interior and wherein the second end includes a tip having aset of spray openings formed therein and wherein the set of sprayopenings are configured to allow hot bleed air to spray into the set offluid passageways such that the hot bleed air is dispersed within theset of fluid passageways and heats the skin.
 14. The air temperaturesensor of claim 13 wherein at least a portion of the skin forms anairfoil.
 15. The air temperature sensor of claim 14 where the set offluid passageways is configured to disperse the hot bleed air to atleast two separate portions of the skin.
 16. The air temperature sensorof claim 15, further comprising a sheath at least partiallycircumscribing the temperature sensor and the sheath shields thetemperature sensor from heat of the at least two separate portions ofthe skin.
 17. The air temperature sensor of claim 15 wherein there areat least two fluid passageways within the interior configured todisperse the hot bleed air to the at least two separate portions of theskin on opposite sides of the airflow path.
 18. The air temperaturesensor of claim 13 wherein at least one of the set of spray openings isconfigured to heat the skin adjacent the tip.
 19. An air temperaturesensor, comprising: a housing having a skin and defining an interior andhaving a distal end that defines an airfoil cross section; a temperaturesensor extending through a portion of the housing and at least partiallyadjacent a portion of the skin; a set of fluid passageways definedwithin the interior and configured to receive hot bleed air and dispersethe hot bleed air to at least two separate portions of the skin; and apiccolo tube having a first end fluidly coupled to receive bleed airfrom a portion of an aircraft engine and a second end, fluidly coupledto the first end, and where the second end includes a tip having a setof spray openings located about the tip and configured to allow bleedair to spray out against specific portions of an inside surface of thehousing.
 20. The air temperature sensor of claim 19 wherein the set ofspray openings are configured to heat at least two separate portions ofthe skin that are wetted surfaces.